Gas turbine engines, such as those utilized in commercial aircraft, include a compressor section that compresses air and a combustor section that ignites combustion gasses mixed with the compressed air. The gasses generated by the combustion section are super-heated and expelled through a turbine section, driving the turbine section to rotate. Absent some form of cooling, the high temperatures of the expelled gasses can cause thermal degradation to occur in the turbine section.
To mitigate thermal degradation from the extreme temperatures, some or all of the turbine stages are actively cooled by passing relatively cool air through the turbine stage. The active cooling increases the life span of the components in the actively cooled turbine stage by reducing breakage resulting from thermal wear. In some example gas turbine engines, the relatively cool air is drawn from a bleed located in the compressor section (referred to as a compressor bleed) and is piped directly to the actively cooled turbine section through a tangential on board injection (TOBI) cooling system.
In practical gas turbine engine systems, air passing through a primary flow path of the turbine section is significantly hotter than air provided to the actively cooled region. Furthermore, in existing gas turbine engines a portion of the air in the primary flow path leaks into a cooling region radially inward of the primary flow path. As a result of the leakage, a temperature at a radially outermost edge of the cooling region is significantly hotter than a temperature at the radially innermost region resulting in a thermal gradient across the turbine stage.
Thermal gradients, such as those caused by the above described leakage, increase the stress on the components experiencing the thermal gradient and can lead to premature wear and/or breakage of the component.